A gas turbine will be explained with reference to FIG. 16. In general, a gas turbine is equipped with a plurality of stages of stationary blades 2 and 3 arrayed in a circle on a casing (a blade circle or a vehicle chamber) 1, and a plurality of moving blades 5 arrayed in a circle on a rotor (a hub of a base) 4. FIG. 16 shows the moving blade 5 at a certain stage, the stationary blade 2 at the same stage (the inlet side of combustion gas 6) as this moving blade 5, and the stationary blade 3 at the next stage (the outlet side of the combustion gas 6) of this moving blade 5.
When pressure loss is large in the gas turbine, turbine efficiency is lowered. Therefore, it is important to improve the turbine efficiency by minimizing the pressure loss.
However, as shown in FIG. 16, there is a case where the moving blade 5 at a certain stage is what is called a free-standing moving blade that has a clearance 8 between a tip 7 of this moving blade 5 and the casing 1. In the case of this free-standing moving blade 5, there is the following problem.
Namely, as shown in FIG. 17, a main flow (shown by a solid-line arrow mark in FIG. 17) of combustion gas 6 flows to the next-stage stationary blade 3 side by passing through between the moving blade 5 and the moving blade 5. In the mean time, in the clearance 8 between the tip 7 of the moving blade 5 and the casing 1, there is generated a leakage flow 9 (shown by a broken-line arrow mark in FIG. 17) that is separate from the main flow of the combustion gas 6.
A mechanism of generating the leakage flow 9 is that as the pressure at a belly surface 10 side of the moving blade 5 is higher than the pressure at a rear surface 11 side of the moving blade 5, the leakage flow 9 is generated from the belly surface 10 side to the rear surface 11 side based on a difference between these pressures.
As shown in FIG. 17, the leakage flow 9 flows at an incidence angle ic to the rear surface 13 side at a front edge 12 of the tip of the stationary blade 3 at the next stage. This leakage flow 9 becomes a flow opposite to the main flow of the combustion gas 6 that flows to the belly surface 14 side of the stationary blade 3.
Therefore, a vortex flow 15 (shown by a solid-line spiral arrow mark in FIG. 17) is generated at the belly surface 14 side of the front edge 12 of the tip of the stationary blade 3. When this vortex flow 15 is generated, pressure loss occurs. The main flow of the combustion gas 6 may deviate from the belly surface 14 side of the stationary blade 3. In FIG. 17, a reference symbol βc denotes an entrance metal angle at the tip portion of the stationary blade 3. Similarly, a reference symbol θc denotes a front-edge including angle at the tip portion of the stationary blade 3. Similarly, a reference number 22 denotes a camber line for connecting between the front edge 12 of the tip portion of the stationary blade 3 and a rear edge 23 of the tip portion.
The incidence angle ic of the leakage flow 9 and the pressure loss have a relative relationship as shown by a solid-line curve in FIG. 18. The solid-line curve in FIG. 18 shows a case of the front-edge including angle θc at the tip portion of the stationary blade 3 shown in FIG. 17.
In this case, the front-edge including angle θc at the tip portion of the stationary blade 3 has been set such that the pressure loss becomes minimum (refer to a point P1 in FIG. 18). However, as described above, the leakage flow 9 is generated, and the pressure loss also becomes large when the incidence angle ic of this leakage flow 9 is large (refer to a point P2 in FIG. 18). When this pressure loss is large, the turbine efficiency is lowered by that amount.
Further, as shown in FIG. 16, seal-air 16 (shown by a two-dot chained line arrow mark in FIG. 16) flows from the rotor 4 side at the upstream of the moving blade 5 at a certain stage. When this seal-air 16 is flowing, there is the following problem.
Namely, the seal-air 16 simply flows out straight in a direction of the height (a radial direction of the turbine) of the moving blade 5 without being squeezed by a nozzle or the like. On the other hand, the moving blade 5 is rotating in a direction of an outline arrow mark together with the rotor 4. Therefore, from the relative relationship between the flow-out of the seal-air 16 and the rotation of the moving blade 5, the seal-air 16 flows at the incidence angle is to the rear-surface side 11 at the front edge 17 of the hub portion of the moving blade 5, as shown in FIG. 17.
As explained above, when the incidence angle is of the seal-air 16 becomes large at the front edge 17 of the hub portion of the moving blade 5 as well, the pressure loss becomes large and the turbine efficiency is lowered by that amount as shown in FIG. 17 and FIG. 18, in a similar manner to that at the front edge 12 of the tip portion of the stationary blade 3.
This problem of the hub portion of the moving blade 5 also applies to a shrouded moving blade in addition to the above-described free-standing moving blade. In FIG. 17, a reference symbol βs denotes an entrance metal angle at the hub portion of the moving blade 5. Similarly, a reference symbol θs denotes a front-edge including angle at the hub portion of the moving blade 5. Similarly, a reference number 24 denotes a camber line for connecting between the front edge 17 of the hub portion of the moving blade 5 and a rear edge 25 of the hub portion.
Further, when the moving blade 5 at a certain stage is a free-standing moving blade, there is the following problem.
Namely, as shown in FIG. 17, the leakage flow 9 is generated from the belly surface 10 side of the moving blade 5 to the rear surface 11 side, at the clearance 8 between the tip 7 of the free-standing moving blade 5 and the casing 1.
Then, as shown in FIG. 19B, a design Mach number distribution shown by a solid-line curve becomes an actual Mach number distribution as shown by a broken-line curve. As a result, on the rear surface 11 of the tip portion 18 of the moving blade 5, deceleration from an intermediate portion to a rear edge 19 is larger in actual Mach distribution G2 than in design Mach distribution G1.
When the deceleration is large, as shown in FIG. 19A, a boundary layer (a portion provided with shaded lines) 20 at a portion from the intermediate portion to the rear edge 19 swells on the rear surface 11 of the tip portion 18 of the moving blade 5. As a result, the pressure loss becomes large, and the turbine efficiency is lowered by that amount. A reference number 21 in FIG. 19 denotes a front edge of the tip portion 18 of the moving blade 5.